Methods and preforms for forming composite members with interlayers formed of nonwoven, continuous materials

ABSTRACT

Materials and methods are provided for producing preform materials for impact-resistant composite materials suitable for liquid molding. Interlayers formed of nonwoven, continuous fibers, such as spunbonded, spunlaced, or mesh fabric, are introduced between non-crimped layers of unidirectional reinforcing fibers to produce a preform for use in liquid-molding processes to produce a composite member. Curing of the preform provides increased impact resistance by increasing the amount of energy required to propagate localized fractures due to impact.

CROSS-REFERENCES TO RELATED APPLICATIONS

This application is a divisional of application Ser. No. 10/974,426,filed on Oct. 27, 2004, which is a continuation-in-part of copendingU.S. patent application Ser. No. 10/428,500, filed on May 2, 2003, andcopending U.S. patent application Ser. No. 10/852,713, filed on May 24,2004, each of which is hereby incorporated herein in its entirety byreference.

BACKGROUND

1. Field

The present invention relates to cured composites built from layers ofunidirectional fibers. In particular, the invention utilizes highlyporous lightweight materials in conjunction with multilayer preforms toobtain cured articles with improved toughness.

2. Description of Related Art

High-performance composite materials built of alternating layers ofunidirectional reinforcing fibers have an advantageous combination ofhigh strength and light weight. As such they find use in aerospace andother industries where such properties are critical. Generally, thecomposite materials are prepared by laying up a number of alternatinglayers wherein adjacent layers have unidirectional fibers running atdifferent angles. The net effect of buildup of several layers of suchunidirectional fabrics is to provide a composite material havingexceptional strength, either quasi-isotropically, or in one or moreparticular directions.

Such composite materials may be produced as prepregs or as preforms. Inprepregs, layers of unidirectional fabrics immersed or impregnated witha matrix material such as a resin are laid-up into the shape of the partto be produced from the composite material. Thereafter, the laid-up partis heated to cure the matrix material and provide the finished compositepart. In the preform approach, layers of unidirectional reinforcingfibers or woven, braided, or warp-knit fabric are laid up similarly tothe way they are laid-up in the prepreg method. However, in the preformmethod, the layers are laid-up dry, i.e., without the matrix material.Thereafter, the laid-up material is infused with the matrix material ina liquid-molding process, and the molded part is heated to cure thematrix material as in the prepreg method.

The alternating layers, or lamina, of reinforcing fibers provide thecomposite articles made from the prepreg or preform process with a greatdeal of strength, especially in directions that align with specificfiber directions. Accordingly, very strong lightweight parts may beproduced, for example, as wings and fuselages of aircraft. Although thealternating lamina of reinforcing fibers provide strength, toughness orimpact resistance is determined mainly by the properties of the curedmatrix material. Impact-resistant or toughened matrix materials aregenerally preferred because they are resistant to damage from impact.This is important, for example, in the case of airplane wings made fromsuch composite materials to avoid failure from foreign-object impactduring flight, damage resulting from ground-maintenance impact (e.g.from tool drop, forklifts, or other vehicles), and the like.Furthermore, because impact damage in composite materials is generallynot visible to the naked eye, it is important for such primaryload-bearing structures to be able to carry their full design load afterimpact and prior to detection using non-destructive techniques.

In prepregs, the matrix material, which is typically an epoxy-basedresin formulation, may be toughened by adding particles of athermoplastic material to the conventional resin. These thermoplasticparticles may either be soluble in the matrix resin and dissolve in theepoxy resin or may be insoluble and placed, during the prepreggingoperation (see, for example, U.S. Pat. No. 5,028,478) on the surface ofeach layer. Upon cure, the thermoplastic resin in the cured epoxy matrixserves to limit crack propagation through the part. Preform materialsmay also be stitched before resin infusion and cure to provide toughnessand crack resistance. However, one drawback to stitching is thereduction of in-plane mechanical properties, particularly as the stitchdensity increases. The prepreg approach of applying particles ofthermoplastic material in the resin before cure is not directlytransferable to the liquid molding processes used to prepare preformarticles. In the resin infusion of the liquid molding process, solublethermoplastics tend to increase the melt-flow viscosity of the matrixresin unacceptably, while insoluble thermoplastic toughening particlestend to be filtered by the preform and thus will not be locateduniformly between the plies in the preform.

In the European Patent EP 1 175 998 to Mitsubishi, laminated productsformed of reinforcing fibers are provided in which thermoplastic resinlayers are provided between layers of the reinforcement fiber. Thethermoplastic resin layer is described in the form of a porous film,fiber, network structure, knitted loop, and the like. The laminatedproduct uses a thermoplastic layer of sufficient permeability betweenthe layers of reinforcing fibers so as not to inhibit liquid resin flowduring a liquid molding process. One drawback inherent in processes suchas those described in EP 1 175 998 is that the preform made ofalternating layers of reinforcing fibers and thermoplastic resin layersare less than perfectly stable during resin infusion. As a result, thereinforcing fibers and the thermoplastic resin layer tend to move orshift during the liquid molding process. Such moving or shifting can bemitigated by stitching together the layers before infusion with theresin. Another drawback to the processes described in EP 1 175 998 isthat they are primarily effective for hand lay-up operations and not forautomated lay-up operations that would be more relevant in thefabrication of large aircraft parts or in the continuous production ofbroad goods.

It would be desirable to provide a molded article made by a preformprocess in which the reinforcing fibers are held tightly in relativeorientation to one another. It would further be desirable to provide aprocess for making such a preform article in widths and lengths feasiblefor producing large-scale parts, such as airplane wings, from them.

SUMMARY

In one embodiment, the invention provides a preform and a compositemember made up of reinforcing layers of unidirectional fibers. Non-woveninterlayers made of spunbonded, spunlaced, or mesh fabric of fibers aredisposed between and stitched to the reinforcing layers. The preform canbe used in a liquid-molding process by which a matrix material isinfused into the preform, followed by heating to gel and set the matrixmaterial. The interlayers are permeable to permit the flow of matrixmaterial during the liquid-molding operation. The layers are securedtogether with stitches or knit threads, such that the unidirectionalfibers are held in place during the infusion process and subsequentcuring of the matrix material to produce a fiber reinforced compositematerial, which can be formed in a desired contoured shape of acomposite member. The material making up the interlayers can increasethe toughness or impact resistance of the finished composite member ascompared to a corresponding member that is formed without theinterlayers or with other materials provided between the layers ofreinforcing fibers. For example, the material of the interlayers canincrease the Mode I or Mode II impact resistance, and the material canbe chosen for compatibility with the matrix material upon curing. In oneembodiment, the matrix material is an epoxy resin and the interlayerfibers are made of a polyamide, polyimide, polyamide-imide, polyester,polybutadiene, polyurethane, polypropylene, polyetherimide, polysulfone,polyethersulfone, polyphenylsulfone, polyphenylene sulfide,polyetherketone, polyethertherketone, polyarylamide, polyketone,polyphthalamide, polyphenylenether, polybutylene terephthalate,polyethylene terephthalate, polyester-polyarylate (e.g. Vectran®),polyaramid (e.g. Kevlar®), polybenzoxazole (e.g. Zylon®), viscose (e.g.Rayon®), carbon-fiber, and glass-fiber.

In another embodiment, the invention provides a method for forming acomposite material. The method includes alternately disposingreinforcing layers formed of fibers of a reinforcing material andinterlayers, which are formed of a nonwoven fabric of continuous fibers.For example, the nonwoven fabric can be formed by spunbonding,spunlacing, or fabric meshing. The material of the interlayers isadapted to increase the impact resistance of the composite material. Insome cases, the interlayers can be formed of a substantially tacklessmaterial so that the reinforcing layers and the interlayers aresubstantially unbonded prior to stitching. The reinforcing layers andthe interlayers are stitched, and the reinforcing layers are infusedwith a matrix material that flows through the interlayers. A compositemember can be formed from the preform by curing the matrix material,with the preform supported in a configuration corresponding to thedesired contour of the finished member.

Fiber-reinforced composite materials may be made by molding a preformand infusing the preform with a thermosetting resin in a number ofliquid-molding processes. Liquid-molding processes that may be used inthe invention include, without limitation, vacuum-assisted resintransfer molding (VARTM), in which resin is infused into the preformusing a vacuum-generated pressure differential. Another method is resintransfer molding (RTM), wherein resin is infused under pressure into thepreform in a closed mold. A third method is resin film infusion (RFI),wherein a semi-solid resin is placed underneath or on top of thepreform, appropriate tooling is located on the part, the part is baggedand then placed in an autoclave to melt and infuse the resin into thepreform. The RFI method is described in U.S. Pat. No. 4,311,661, titled“Resin Impregnation Process,” which issued Jan. 19, 1982, the entirecontents of which is incorporated herein by reference.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention will become more fully understood from thedetailed description and the accompanying drawings, wherein:

FIGS. 1 a-1 d are section views schematically illustrating fibers forthe interlayers of preforms and composite members formed according tovarious embodiments of the present invention;

FIG. 2 is a section view illustrating a stitched multiaxial fabric of apreform according to one embodiment of the present invention;

FIG. 3 is a perspective view illustrating a stitched preform for forminga composite member according to yet another embodiment of the presentinvention;

FIG. 4 is an elevation view illustrating a composite member formed withthe preform of FIG. 3 using a contoured mold;

FIG. 5 is an elevation view schematically illustrating a process forpreparing a multiaxial fabric with multiple layers and interlayersaccording to another embodiment of the present invention; and

FIG. 6 is a perspective view schematically illustrating the fabricformed according to the process of FIG. 5.

DETAILED DESCRIPTION

The present invention now will be described more fully with reference tothe accompanying drawings, in which some, but not all embodiments of theinvention are shown. This invention may be embodied in many differentforms and should not be construed as limited to the embodiments setforth; rather, these embodiments are provided so that this disclosurewill be thorough and complete, and will fully convey the scope of theinvention to those skilled in the art. Like numbers refer to likeelements throughout.

FIG. 2 illustrates a multiaxial fabric 10 that is made of alternatinglayers 2 of reinforcing fibers and interlayers 6 according to oneembodiment of the present invention. Each interlayer 6 is typically anonwoven configuration, such as a spunbonded, spunlaced, or mesh fabricof thermoplastic fibers. The interlayers 6 are disposed between andknit-stitched to the reinforcing layers 2. The interlayers 6 need not beotherwise bonded or connected to the layers 2 of the reinforcing fibers.Fabrics in which the interlayers 6 and reinforcing layers 2 aremelt-bonded together are further described in U.S. application Ser. No.10/428,500, titled “Highly Porous Interlayers to Toughen Liquid-MoldedFabric Based Composites,” filed May 2, 2003. In either case, theresulting multiaxial fabric 10 may be manufactured by a number ofprocesses to produce preforms that are 12-300″ wide, and typically atleast about 50″ wide. For example, if the interlayers 6 are formed witha nonwoven configuration, e.g., a fabric that is spunbonded, spunlaced,or mesh fabric, each interlayer 6 can be formed by an automated methodand with relatively wide widths that can be difficult or impossible toform by braiding, weaving, and the like. The fabric 10 can be used as apreform 20 (FIG. 3) that is used to form a composite member 22 (FIG. 4)having a desired configuration, e.g., by disposing the preform 20 in amold 24, infusing the preform 20 with a matrix material such as athermosetting resin in a liquid-molding process, and then heating thepreform 20 in the mold to gel and set the matrix material. Theinterlayers 6 can be lightweight and porous to thereby minimizedistortion of the layers 2 of the reinforcing fibers and reduce theresistance of a flow of the matrix material through the interlayers 6during infusion of the layers 2 of the reinforcing material.

Each layer 2 of the reinforcing fibers is typically a layer ofunidirectional fibers. Such layers 2 of unidirectional fibers for use inmultiaxial preforms and fiber reinforced composite materials are wellknown in the art. For example, the unidirectional fibers can be made ofcarbon fibers. Other examples of unidirectional fibers include, withoutlimitation, glass fibers and mineral fibers. The layers 2 ofunidirectional fibers can be prepared by a laminating process in whichunidirectional carbon fibers are taken from a creel containing multiplespools of fiber that are spread to the desired width and layered with ainterlayer 6.

Each interlayer 6 is typically made of a spunbonded, spunlaced, or meshfabric of fibers, which can be thermoplastic. The fibers may be selectedfrom among any type of fiber that is compatible with the thermosettingmatrix material used to form the fiber reinforced composite material ofthe member 22. For example, the fibers of the interlayer 6 may beselected from the group consisting of polyamide, polyimide,polyamide-imide, polyester, polybutadiene, polyurethane, polypropylene,polyetherimide, polysulfone, polyethersulfone, polyphenylsulfone,polyphenylene sulfide, polyetherketone, polyethertherketone,polyarylamide, polyketone, polyphthalamide, polyphenylenether,polybutylene terephthalate, polyethylene terephthalate,polyester-polyarylate (e.g. Vectran®), polyaramid (e.g. Kevlar®),polybenzoxazole (e.g. Zylon®), viscose (e.g. Rayon®), carbon-fiber, andglass-fiber.

Generally, the interlayers 6 are formed of any of various thermoplasticmaterials that are chemically compatible with the matrix resin and thatdo not dissolve during infusion and cure into the matrix. The materialof the interlayers 6 is not soluble to any great extent in theunderlying matrix, except as to facilitate better contact and/oradhesion between the interlayer and the matrix. The melting point of thethermoplastic material of the interlayers 6 is typically near or abovethe cure temperature of the matrix resin to ensure that compositeproperties, such as elevated-temperature compression strength, are notcompromised. The thermoplastic materials also have good resistance tosolvents like ketones, water, jet fuel, and brake fluids to ensure thatthe composite material does not become susceptible to strengthdegradation is exposed to these solvents. Although the present inventionis not limited to any particular theory of operation, it is believedthat, in order for the interlayers 6 to provide a desired increase inthe impact resistance of the resulting composite member 22, thethermoplastic material must have some chemical compatibility with theresin (e.g. chemical bonding, hydrogen bonding, etc.); the thermoplasticmaterial must have sufficient inherent toughness (i.e., not toobrittle); and the thermoplastic material must have a fairly high modulus(i.e., not so low that the thermoplastic material acts like aplasticizing layer and reduces properties).

In one embodiment, the fibers of the interlayers 6 are made from two ormore materials. For example, the two or more materials may be preparedby mechanically mixing different fibers, which are used to create thespunbonded, spunlaced, or mesh fabric. Alternatively, the two or morematerials may be used to form a bi-component fiber, tri-component fiber,or higher component fiber for use as the interlayer 6. That is, eachfiber of the fabric used for the interlayer 6 can include multiplematerials. Non-limiting examples of multi-component fibers areillustrated schematically in FIG. 1. FIG. 1( a) shows in cross-section afiber made for example by coextrusion of a fiber material A and a fibermaterial B. Such a fiber may be produced by a spinneret with twooutlets. FIG. 1( b) shows a bi-component fiber made from materials A andB such as would be produced by extrusion through four spinnerets.Similarly, FIG. 1( c) shows a bi-component fiber spun from eightspinnerets. In another embodiment, the bi-component fiber is provided inthe form of a core sheath fiber such as illustrated in FIG. 1( d). In acore-sheath fiber, a fiber material of one type, illustrated as B inFIG. 1( d), is extruded as the core, while a fiber material of anothertype, illustrated as A in FIG. 1( d) is extruded as the sheath. Forexample, the fiber material A of the sheath can be a polyurethane, andthe fiber material B of the core can be a polyamide.

Bi-component fibers such as those illustrated in FIG. 1 and other fiberscontaining more than two components are well known in the art and can bemade by a number of conventional procedures. Additionally, although thefibers in FIG. 1 are illustrated schematically with circularcross-sections, it is to be appreciated that other cross-sections may beused.

Each fiber of the interlayers 6 can have any size, e.g., according tothe particular application for the interlayer 6 and the resultingcomposite member 22. In particular, the fibers making up the interlayer6 can have diameters from 1 to 100 microns, e.g., from 10 to 75 microns,such as from 10 to 30 microns. In another embodiment, the fibers havediameters from 1 to 15 microns.

The material used for the interlayers 6 can also have a wide range ofareal densities. The areal density may be chosen according to the amountrequired to impart the desired impact resistance, as verified forexample by compression-after-impact testing according to Boeing testmethod BSS 7260 (also known as SACMA SRM 2-88.) The desiredimpact-resistance level is determined on a part-by-part basis assumingspecific impact-energy levels. In one embodiment, the interlayermaterial has an areal density of 1-50 grams/square meter. In anotherembodiment, the areal density of each interlayer 6 is about 2-15grams/square meter, such as between about 5 and 15 grams/square meter.The optimum areal weight for any particular composite member 22 can bedetermined by trial and error, but typically is between about 2% and 4%of the overall composite weight.

The interlayer material may be a spunbonded fabric. Spunbonded fabricsare produced from continuous fibers that are continuously spun andbonded thermally. These fabrics are commercially available from a widevariety of sources, primarily for the clothing industry. The spunbondedfabrics for use in the present invention typically have areal weightsthat are generally lower than those of fabrics used in clothing.

In another embodiment, each interlayer 6 is a spunlaced fabric.Spunlaced fabrics are prepared from continuous fibers that arecontinuously spun and bonded mechanically. These fabrics arecommercially available from a wide variety of sources, primarily for theclothing industry. Spunlaced fabrics for use in the present inventiontypically have areal weights that are generally lower than thosecommonly used in the clothing industry.

In yet another embodiment, each interlayer 6 comprises a nonwoven meshfabric. For example, the mesh construction of the interlayer 6 cancontain between about 0.5 and 15 threads or fibers per inch in the warpand weft directions.

As shown in FIGS. 2 and 3, the multiaxial preform 20 comprises aplurality of the reinforcing layers 2 with the interlayers 6 disposedbetween the reinforcing layers 2 and stitched to attach and stabililzethe reinforcing layers 2 and interlayers 6 into a multiaxial fabric.Typically, each multiaxial preform 20 has 4 or more reinforcing layers2, but the preform 20 can have fewer layers 2. For example, the preform20 can have between 2 and 16 layers (or lamina) of the reinforcinglayers 2.

The lamina of unidirectional fibers in the multiaxial fabric 10 may belaid-down in quasi-isotropic or orthotropic patterns. The pattern may berepeated as needed to achieve a desired thickness of the finishedcomposite member 22. The repeated pattern may be constant, or may bevaried across the preform 20. Where the repeated pattern is variedacross the preform 20, the locally different thicknesses may bemechanically held in place, such as by stitching the lamina of thelayers 2 and interlayers 6 together.

For example, in some cases, the lamina of the reinforcing layers 2 arelaid-down in a quasi-isotropic pattern. A quasi-isotropic pattern is onethat approximates an isotropic material in the plane of the fibers. Thisis also known as transverse isotropy. For example, FIGS. 5 and 6illustrate a +45/−45 pattern, i.e., in which a first layer 2 a isdisposed at +45° and a second layer 2 b is disposed at a −45° relativeto the transverse direction of the fabric 10. It is also possible todispose lamina in a quasi-isotropic 0/+45/90/−45 pattern. Otherquasi-isotropic patterns include +45/0/−45/−90; −45/0/+45/90; and0/+60/−60.

Alternatively, the lamina of the reinforcing layers 2 may be laid-downin an orthotropic pattern. Orthotropic means having fibers or units suchthat the net result is not quasi-isotropic in plane like thequasi-isotropic patterns just described. An example of an orthotropicpattern is one with 44% 0°, 22% +45°, 22% −45° and 12% 90° fibers. Inthis example, greater longitudinal strength (along the 0°-direction) andlower shear strength (±45°-direction) and transverse strength(90°-direction) than a quasi-isotropic (25/50/25) lay-up are achieved.The resulting built-up lamina provide higher strength and thickness inthe 0° direction as compared to a quasi-isotropic laminate, but providelower shear strength and thickness (provided by the ±45° layers 2).Correspondingly, in this example, the 90° strength is lower than aquasi-tropic laminate. The term orthotropic is well understood in thefield. For example, a 0° fabric is orthotropic, as well as any otherpattern that does not result in balanced average in plane (i.e.quasi-isotropic) properties.

As noted above, it is common to construct a fabric such that itcomprises a set of four laminae. The multilayer fabric layer is commonlyreferred to as a stack. Where desired, a preform construction willcomprise a pattern of the four laminae to achieve a desired thickness.For example, when it is desired to build-up a desired thickness,mirror-image lamina stacks are used to prevent post-cure bending andtwisting due to thermal stresses created after curing the resin atelevated temperature. In such a case, the total lay-up would be made upof groups of balanced laminae, or laid-up alternately to balance thelaminate. This practice is common in the field and is done to ensure thefabrication of flat parts and to avoid the problem of parts with unknownand/or temperature-sensitive configurations.

As previously noted, the interlayers 6 can be stitched (i.e., knitted orsewed) with the thread 8 to the unidirectional fiber layers 2 tomaintain the orientation of the layers 2 in place during resin infusioninto the mold 24 during a (subsequent) liquid-molding process. Forexample, each stitch of the thread (referred to collectively byreference numeral 8) can connect one or more of the interlayers 6 to oneor more of the layers 2. Thus, a warp-knit, multiaxial fabric 10 may beassembled by knit-stitching the reinforcing layers 2 together withinterlayers 6 between the reinforcing layers 2. The knit or sewingthread 8 may be selected from a variety of materials, including withoutlimitation, polyester-polyarylate (e.g. Vectran®), polyaramid (e.g.Kevlar®), polybenzoxazole (e.g. Zylon®), viscose (e.g. Rayon®), acrylic,polyamide, carbon, and fiberglass. Where desired, the knitting or sewingstep is carried out after the initial lay-up of the multiaxial preform20. The same kinds of threads 8 may be used to hold locally differentthicknesses mechanically in place by stitching and by tufting.

FIG. 2 shows the fabric 10 of the multiaxial preform 20 for use in aliquid-molding process to form the composite member 22. In FIG. 2,interlayers 6 made of thermoplastic fibers are disposed betweenreinforcing fabric layers 2 of unidirectional fabrics. Multiple sewingthreads 8 are used to hold the lamina 2, 6 of the preform 20 together.As shown in FIG. 2, each thread 8 (and each stitch) extends through eachof the layers 2 and interlayers 6 of the preform 20 in alternatedirections.

Thus, all of the layers 2 and interlayers 6 can be connected bystitching, with none of the layers 2 and interlayers 6 being melt-bondedor otherwise bonded together. In this regard, the material of theinterlayer 6, in some cases, can provide little or no tackiness orstickiness for bonding or adhering to the layers 2. Instead, thestitches 8 can provide any necessary connection between the layers 2 andinterlayers 6 and/or mechanical fasteners can be provided for temporaryor permanent connection of the layers 2 and interlayers 6. Thus, thefabric 10, the preforms 20, and/or the composite members 22 can beformed according to the present invention without the use of“tackifiers,” i.e., materials for bonding the layers 2 and interlayers6. That is, the stitches 8 can connect the layers 2 and interlayers 6during the final stacking process and during the infusion process. Thelack of a tackifier between the layers 2 and interlayers 6 can increasethe penetrability of the preform 20 by the matrix material and therebyfacilitate infusion by the matrix material.

Further, the interlayers 6 can be formed of materials that improvespecific characteristics of the resulting members 22, such as the impactresistance or toughness, regardless of the tackiness of the interlayermaterial. Thus, the impact resistance of the resulting composite member22 can be greater in some cases than the impact resistance that can beachieved in a member of similar dimensions that is formed withinterlayers 6 of other materials, such as those materials that providetackiness, or the impact resistance that can be achieved in a member ofsimilar dimensions that is formed with the composite layers disposedadjacently with no interlayers therebetween. The impact resistance ofthe composite fabric 10 or member 22 can generally be increased byincreasing the amount of energy required to propagate localizedfractures due to impact.

A device such as described in U.S. Pat. No. 5,241,842 to Hagel, or U.S.Pat. No. 6,276,174 to Wunner, et al. (the disclosures of which areincorporated by reference) may be used to prepare multiaxial preforms byproviding tows of unidirectional carbon fibers. One or a plurality oftows is pulled across pins to create reinforcing layers ofunidirectional fibers. In this embodiment, a means is provided forintroducing the interlayer material between the layers 2 ofunidirectional carbon fibers. Because the interlayer material isnon-directional, it need not be introduced at an angle in the way thatthe unidirectional carbon fibers are.

FIG. 5 illustrates the process of forming the multiaxial fabric 10according to one embodiment of the present invention. As illustrated,two layers, referred to individually by reference numerals 2 a, 2 b arestacked with one interlayer 6 therebetween, though in other embodimentsmore layers 2 and interlayers 6 can be provided. In some cases,interlayers of the thermoplastic material, referred to individually byreference numerals 6 a, 6 b, can also be provided on the top and/orbottom of the fabric 10 as shown. Any number of the interlayers can beprovided. For example, referring to the interlays 6, 6 a, 6 billustrated in FIGS. 5 and 6, the fabric 10 can include only theinterlayer 6 between the layers 2 a, or the interlayer 6 in combinationwith one or both of the interlayers 6 a, 6 b. Further, each of theinterlayers 6 can have different thicknesses. For example, theinterlayers 6 a, 6 b on the top and bottom of the fabric 10 can bethinner, such as about half as thick, as the interlayer 6 disposedbetween the layers 2 a, 2 b. In the illustrated embodiment, the firstlayer 2 a is a +45 layer of carbon and the second layer 2 b is a −45layer of carbon. A perspective view of the fabric 10 is shown in FIG. 6.The layers 2 a, 2 b and interlayers 6, 6 a, 6 b can pass through a nipbetween rollers 17, and the fabric 10 is then processed by a knittingunit 56 that forms stitches 8 in the fabric 10, as shown in FIG. 6.

The multiaxial preforms of the invention may be made into curedfiber-reinforced composite materials by a variety of liquid-moldingprocesses. In one, vacuum-assisted resin transfer molding, a matrixmaterial such as a resin is introduced to a mold containing themultiaxial preform under vacuum. As illustrated in FIG. 4, the mold 24typically defines one or more surfaces corresponding to a desiredcontour of the finished composite member 22 so that the preform 20 issupported in the desired configuration. The matrix material infuses thepreform 20 and saturates the interlayers 6 between the layers 2 ofunidirectional fibers. The interlayers 6 are made of a material that ispermeable to permit the flow of the matrix material during theliquid-molding operation. Furthermore, the stitches 8 between the layers2 and interlayers 6 hold the unidirectional fibers in place during theinfusion process. The mold 24, which is only partially illustrated inFIG. 4, can be a closed vessel-like device for containing the vacuum. Inanother method, typically referred to as resin transfer molding, resinis infused under pressure into a closed mold. These and otherliquid-molding processes may be used to prepare the curedfiber-reinforced composite material of the invention.

Following infusion of the resin in the mold 24 in a process such asthose described above, the mold 24 is heated to cure the resin toproduce the finished composite member 22. During heating, the resin orother matrix material reacts with itself to form crosslinks in thematrix of the composite material. After an initial period of heating,the resin gels. At gel, the resin no longer flows, but rather behaves asa solid. In some cases, the resin can be gelled at a temperature that isbelow the melting point of the thermoplastic fibers of the interlayer 6in order to prevent their melting and flowing into the reinforcementfiber bundles. After gel, the temperature may be ramped up to a finaltemperature to complete the cure. The final cure temperature depends onthe nature and properties of the thermosetting resin chosen. For thecase of aerospace-grade epoxy resins, it is conventional to ramp thetemperature after gel up to a temperature range of 325 to 375° F. andhold at this temperature for 1 to 6 hours to complete the cure.

For example, it is known that, in certain instances, thermoplasticdissolved into a matrix resin may increase a resin's Mode I fracturetoughness. In a composite subjected to impact, the increase in a resin'sMode I fracture toughness has the effect of reducing the amount ofmatrix cracking that occurs due to impact. Alternatively, this may bestated that the increase in the resin's Mode I fracture toughnessrequires more energy to form matrix cracks in the composite due to animpact than an otherwise equivalent composite made with an untoughenedversion of the same matrix resin. Thus, the increase in the matrix ModeI fracture toughness reduces the amount of damage due to an impact,specifically by reducing the amount of cracking in the matrix.

Undissolved layers, if selected appropriately, may increase the Mode IIfracture toughness of the composite. In general, they do little, ifanything, to change the fracture toughness of the matrix resin. In acomposite subjected to impact, the increased Mode II fracture toughnessis manifested by a reduced propensity for delaminations to occur betweenthe plies of the composite material. Alternatively, this may be statedthat the increase in the Mode II fracture toughness of the compositerequires more energy to delaminate the composite or for damage to growlaterally (perpendicular to the direction of the impact and in the planeof the structural fibers) in the composite than an otherwise equivalentuntoughened composite. Thus, the increase in the composite's Mode IIfracture toughness reduces the amount of damage due to an impact,specifically by reducing the tendency of the damage to grow laterallyaway from the point of impact, thereby constraining the volume ofmaterial that is damaged due to impact and increasing the composite'sresidual strength after impact.

Thus, the fracture toughness of the resulting composite members 22formed according to the present invention can be greater than thefracture toughness of other composite members, such as composite membersformed with interlayers 6 between the layers 2 of the reinforcingmaterial. In some cases, the compression-after impact (CAI) strength canbe increased by approximately 100% and the impact-damage area can bedecreased about 90% for a ½″ impact tup at 270 in-lb impact energy.

In some cases, the preforms 20 and composite members 22 can becharacterized by an increased impact resistance provided by theinterlayers 6 such that the need for stitching is reduced. That is,while some conventional composite members include stitches to increaseimpact resistance, such increases in impact resistance can instead beprovided by the interlayers 6 of the present invention. Thus, accordingto one embodiment of the present invention, the stitching of thecomposite material can be provided primarily for securing the layers 2and interlayers 6 in the desired configuration, and the amount ofstitching can be less than that which would be required for increasingimpact resistance in the absence of the interlayers 6.

EXAMPLES

The results shown below are for compression-after-impact (CAI) panelsmade and tested according to BMS 8-276 (a Boeing material specificationfor a toughened prepreg system used for commercial aircraft) using BSS7260 Type II, Class 1 impact with an impact energy of 270 in-lb.

Test panels were prepared as follows. The panel lay-up was(+45/0/−45/90)_(3S) using unidirectional fabric from AnchorReinforcements (Huntington Beach, Calif.) to which spunbonded fabric hadbeen melt-bonded. A control used only a thermoplastic weft fiber to holdthe fabric together. The three spunbonded fabrics were supplied bySpunfab (Cuyahoga Falls, Ohio) in areal weights of 0.125, 0.250, and0.375 oz/yd². The three materials used were PE2900, a polyester; VI6010,a ternary polymer blend; and PA1008, a polyamide.

A dry, uni-directional tape 13-inches in width was prepared bymelt-bonding the respective spunbonded fabrics onto a tape containing190 g/m² of T700 carbon fibers (Toray, Tokyo, Japan). Theuni-directional tape was cut in the same manner as prepreg and laid-upaccording to BMS 8-276 as described above. The laid-up fabric was VARTMprocessed using an epoxy resin, TV-15, from Applied Poleramic, Inc.(Benicia, Calif.). After infusion and cure, the resulting panels weremachined into 4″×6″ impact test specimens according to BSS7260. Impactwas performed using a 0.3125″ spherical steel tup. Four panels for eachconstruction were tested.

After impact, all specimens were ultrasonically C-scanned. Impact damageareas were calculated directly from the center “hole” shown in theamplitude plots using the built-in software tool on the C-scanapparatus. These results are shown in Table 1.

Compression-after-impact strength results are shown in Table 2 and panelthicknesses and per-ply thicknesses are shown in Table 3. Tables 1 and 2show significant decreases in impact damage area for the PA1008 andVI6010 interlayer materials as well as significant increases incompression-after-impact strength for these same materials,respectively. Table 3 shows that the interlayer-toughening concept meetsthe current commercial Boeing specification (BMS 8-276) for per-plythickness.

TABLE 1 Average Impact Damage Area for Three Panels vs. Control. PercentChange in Impact Impact Damage Area (in²) Area Interlayer Areal WeightInterlayer Areal Weight Spunbonded 0.125 0.250 0.375 0.125 0.250 0.375Examples Fabric none oz/yd² oz/yd² oz/yd² oz/yd² oz/yd² oz/yd²Comparative Control 7.134 N/A N/A N/A 1 PE2900 8.258 8.632 10.037 15.821.0 40.7 2 V16010 4.529 3.936 2.093 −36.5 −44.8 −70.,7 3 PA1008 1.4891.160 0.619 −79.1 −83.7 −91.3

TABLE 2 Average Compression-After-Impact Strength for Three Panels vs.Control. Percent Change CAI Strength (ksi) CAI Strength InterlayerInterlayer Areal Weight Areal Weight Exam- 0.125 0.250 0.375 0.125 0.2500.375 ples none oz/yd² oz/yd² oz/yd² oz/yd² oz/yd² oz/yd² Com- 19.3 par-ative 1 17.2 15.6 13.8 −10.9 −19.2 −28.6 2 20.5 24.3 29.4 6.1 26.2 52.63 30.6 27.8 39.6 58.6 44.4 105.3

TABLE 3 Average Cured-Panel Thicknesses. Average Panel Thickness (in)Average Per-Ply Thickness (mil)* Interlayer Areal Weight InterlayerAreal Weight 0.125 0.250 0.375 0.125 0.250 0.375 Examples none oz/yd²oz/yd² oz/yd² none oz/yd² oz/yd² oz/yd² Comparative 0.170 7.08 1 0.1750.185 0.186 7.29 7.71 7.75 2 0.173 0.182 0.180 7.21 7.58 7.50 3 0.1770.189 0.187 7.38 7.88 7.79 *calculated from average thickness/24 plies

Many modifications and other embodiments of the invention will come tomind to one skilled in the art to which this invention pertains havingthe benefit of the teachings presented in the foregoing descriptions andthe associated drawings. Therefore, it is to be understood that theinvention is not to be limited to the specific embodiments disclosed andthat modifications and other embodiments are intended to be includedwithin the scope of the appended claims. Although specific terms areemployed herein, they are used in a generic and descriptive sense onlyand not for purposes of limitation.

What is claimed is:
 1. A preform for forming a composite member, the preform comprising: a plurality of reinforcing layers formed of fibers of a reinforcing material; a plurality of interlayers, the interlayers being disposed alternately between the reinforcing layers, each interlayer being formed of a nonwoven fabric of continuous thermoplastic fibers, different than the fibers of the reinforcing layers, such that the layers and interlayers are configured to be infused with a matrix material, wherein the plurality of interlayers have different respective thicknesses; and stitching extending through the reinforcing layers and the interlayers and configured to engage the reinforcing layers and the interlayers, wherein a material of the interlayers is adapted to increase the impact resistance of the composite member.
 2. A preform according to claim 1 wherein each interlayer comprises a nonwoven fabric formed by at least one of the group consisting of spunbonding, spunlacing, and fabric meshing.
 3. A preform according to claim 1 wherein the material of the interlayers comprises at least one of the group consisting of polyamide, polyimide, polyamide-imide, polyester, polybutadiene, polyurethane, polypropylene, polyetherimide, polysulfone, polyethersulfone, polyphenylsulfone, polyphenylene sulfide, polyetherketone, polyethertherketone, polyarylamide, polyketone, polyphthalamide, polyphenylenether, polybutylene terephthalate, polyethylene terephthalate, polyester-polyarylate, polyaramid, polybenzoxazole, viscose, carbon-fiber, and glass-fiber.
 4. A preform according to claim 1 wherein the material of the interlayers is substantially tackless.
 5. A preform according to claim 1 wherein the material of the interlayers is adapted to remain intact while the matrix material is infused therethrough and cured.
 6. A preform according to claim 1 wherein a width of the preform is at least about 50 inches.
 7. A preform according to claim 1 wherein the fabric of each interlayer comprises a mechanical mix of dissimilar fibers.
 8. A preform according to claim 1 wherein the fabric of each interlayer comprises multi-component fibers.
 9. A preform according to claim 1 wherein the respective thicknesses of the interlayers are dependent upon a relative position of the interlayers with respect to the plurality of reinforcing layers.
 10. A preform according to claim 1 wherein a first interlayer is positioned between a pair of reinforcing layers and a second interlayer is positioned proximate an outer face of an outermost reinforcing layer, wherein the second interlayer is thinner than the first inter layer.
 11. A composite member defining a predetermined configuration, the composite member comprising: a plurality of reinforcing layers formed of fibers of a reinforcing material; a plurality of interlayers, the interlayers being disposed alternately between the reinforcing layers, each interlayer being formed of a nonwoven fabric of continuous thermoplastic fibers, different than the fibers of the reinforcing material, wherein the plurality of interlayers have different respective thicknesses; stitching extending through the reinforcing layers and the interlayers and configured to engage the reinforcing layers and the interlayers; and a matrix material infused in the reinforcing material, the matrix material being cured with the reinforcing layers and interlayers defining the predetermined configuration, wherein a material of the interlayers is adapted to increase the impact resistance of the composite member.
 12. A composite member according to claim 11 wherein each interlayer comprises a nonwoven fabric formed by at least one of the group consisting of spunbonding, spunlacing, and fabric meshing.
 13. A composite member according to claim 11 wherein the material of the interlayers comprises at least one of the group consisting of polyamide, polyimide, polyamide-imide, polyester, polybutadiene, polyurethane, polypropylene, polyetherimide, polysulfone, polyethersulfone, polyphenylsulfone, polyphenylene sulfide, polyetherketone, polyethertherketone, polyarylamide, polyketone, polyphthalamide, polyphenylenether, polybutylene terephthalate, polyethylene terephthalate, polyester-polyarylate, polyaramid, polybenzoxazole, viscose, carbon-fiber, and glass-fiber.
 14. A composite member according to claim 11 wherein the material of the interlayers is substantially tackless.
 15. A composite member according to claim 11 wherein a width of the composite member is at least about 50 inches.
 16. A composite member according to claim 11 wherein the fabric of each interlayer comprises a mechanical mix of dissimilar fibers.
 17. A composite member according to claim 11 wherein the fabric of each interlayer comprises multi-component fibers.
 18. A composite member according to claim 11 wherein the respective thicknesses of the interlayers are dependent upon a relative position of the interlayers with respect to the plurality of reinforcing layers. 